Traditional metal aircraft structures are fastened with rivets, hi-loks fasteners, or equivalent sheer fasteners. However, composite materials differ from metals in that they have quasi-isentropic properties, which raises two issues in fastening composite structures. The first issue is in drilling the holes. Drilling or punching a hole in a composite structure severs the continuous fibers and effectively isolates the load carrying capability of the fibers between the hole and the edge of the composite structure. This requires the placement of additional edge material to separate the hole from the edge of the composite structure (e.g., an aircraft component, such as a panel), which increases the cost, weight, and manufacturing time of the component. The second issue is due to bearing weakness in the composites. This is caused by the composite's low out-of-plane sheer resistance. Due to this factor, most composites are sized to a critical bearing load, not to a strength load.
While a number of composite washers that focus on sealing and conforming to contour have been disclosed, these composite washers do not add any structural benefit to the composite structure. Similarly, there exist composite fasteners that also focus primarily on sealing the hole, however, they do not transfer load within the composite structure. Moreover, a variety of inserts for potted cores exist, but they are primarily independent fasteners that are added after the manufacturing of the panel and do not use fibers to transfer load into the panel.
There are methods for providing integral reinforcements of sheet metal structures, such as automotive panels. However, these methods are directed at metal, and would cause delamination if used in a composite structure. For example, U.S. Pat. No. 6,969,551 describes a method and assembly for fastening and reinforcing a structural member (e.g., an automotive vehicle pillar, such as an automotive “B” pillar). The assembly preferably includes a structural member having a first and a second portion defining a cavity therebetween. An expandable reinforcement material, such as an epoxy-based reinforcement material, and a spacer may be disposed within the cavity. The assembly also typically includes a fastener or fastening assembly. In operation, the spacer preferably assists in supporting the first and second portions of the structural member during changes of state (e.g., softening) of the reinforcement material.
Other methods address ensuring the integrity of the panel during impact, but are not designed for the purpose of transferring load from a fastener. For example, U.S. Patent Publication No. 2006/0090673 discloses a composite structure having front and back faces, the panel comprising facing reinforcement, backing reinforcement, and matrix material binding to the facing and backing reinforcements, the facing and backing reinforcements each independently comprising one or more reinforcing sheets, the facing reinforcement being located on or embedded in matrix material adjacent to the front face of the panel, the backing reinforcement being located in a plane or planes substantially parallel to the plane or planes of the facing reinforcement, and being substantially coextensive therewith, and spaced therefrom by matrix material, the facing and backing reinforcements being interconnected to resist out-of-plane relative movement. The reinforced composite structure is useful as a barrier element for shielding structures, equipment, and personnel from blast and/or ballistic impact damage.
Other methods address reinforcement of composite joints, focusing on adding reinforcement after layup, but do not involve components integral to the structure. For example, U.S. Patent Publication No. 2010/0320320 discloses a structure for an aerospace vehicle having a composite beam chord clamped between the first and second metal plates. The beam chord is clamped at a force that precludes or reduces beam chord delamination under axial loading during operation of the vehicle. Yet, other methods for fastener reinforcement involve double-plys, but fail to specify any unique components specifically focused on bearing load or edge distance optimization. For example, U.S. Patent Publication No. 2010/0065688 discloses a computer-implemented method and apparatus for creating a structural joint for an aircraft. A first composite component and a second composite component are co-joined to form the structural joint for the aircraft. A hole through the first composite component and the second composite component is created. A composite shear pin is placed through the hole. A composite collar is bonded to the composite shear pin with an adhesive. Finally, other methods utilize a metal spray that wicks into the laminate after a hole is drilled. For example, U.S. Pat. No. 4,763,399 discloses a method of strengthening a bolt hole in a fibrous composite laminate when a hole is drilled in a laminate sheet. The method allows for increased thickness strength, bearing strength, and the use of low-cost, high-strength fasteners, which may or may not be galvanically compatible through the use of plating the edges of the holes with a metal spray.
As set forth above, current methods include adding additional plys around holes and requiring large edge distances. However, such current methods result in an increase in material, and therefore, an increase in weight and cost for each part. Accordingly, a need exists for a system and method to increase the strength of composite materials without increasing the weight or cost for each part.